1. Field of the Invention
The invention relates to abradable surfaces with locally varying porosity or abradability, for turbine engines, including gas or steam turbine engines, the engines incorporating such abradable surfaces, and methods for reducing engine blade tip wear and blade tip leakage. Local porosity or abradability is selectively varied through use of holes or dimple depressions of desired polygonal profiles that are formed into the surface of otherwise monolithic abradable surfaces or rib structures. Abradable porosity within a rib is varied locally by changing any one or more of hole/depression depth, diameter, array pitch density, and/or volume. In various embodiments, localized porosity increases and corresponding abradability increases axially from the upstream or forward axial end of the abradable surface to the downstream or aft end of the surface. In this way, the forward axial end of the abradable surface has less porosity to counter hot working gas erosion of the surface, while the more aft portions of the abradable surface accommodate blade cutting and incursion with lower likelihood of blade tip wear.
2. Description of the Prior Art
As described in the aforementioned U.S. Pat. No. 8,939,707, known turbine engines, including gas turbine engines and steam turbine engines, incorporate shaft-mounted turbine blades circumferentially circumscribed by a turbine casing or housing. Hot gasses flowing past the turbine blades cause blade rotation that converts thermal energy within the hot gasses to mechanical work, which is available for powering rotating machinery, such as an electrical generator. Referring to FIGS. 1-6, known turbine engines, such as the gas turbine engine 80 include a multi-stage compressor section 82, a combustor section 84, a multi-stage turbine section 86 and an exhaust system 88. Atmospheric pressure intake air is drawn into the compressor section 82 generally in the direction of the flow arrows F along the axial length of the turbine engine 80. The intake air is progressively pressurized in the compressor section 82 by rows rotating compressor blades and directed by mating compressor vanes to the combustor section 84, where it is mixed with fuel and ignited. The ignited fuel/air mixture, now under greater pressure and velocity than the original intake air, is directed to the sequential rows R1, R2, etc., in the turbine section 86. The engine's rotor and shaft 90 has a plurality of rows of airfoil cross sectional shaped turbine blades 92 terminating in distal blade tips 94 in the compressor 82 and turbine 86 sections. For convenience and brevity further discussion of turbine blades and abradable layers in the engine will focus on the turbine section 86 embodiments and applications, though similar constructions are applicable for the compressor section 82. Each blade 92 has a concave profile high-pressure side 96 and a convex low-pressure side 98. The high temperature and pressure combustion gas, flowing in the combustion flow direction F imparts rotational motion on the blades 92, spinning the rotor. As is well known, some of the mechanical power imparted on the rotor shaft is available for performing useful work. The combustion gasses are constrained radially distal the rotor by turbine casing 100 and proximal the rotor by air seals 102. Referring to the Row 1 section shown in FIG. 2, respective upstream vanes 104 and downstream vanes 106 direct upstream combustion gas generally parallel to the incident angle of the leading edge of turbine blade 92 and redirect downstream combustion gas exiting the trailing edge of the blade.
The turbine engine 80 turbine casing 100 proximate the blade tips 94 is lined with a ring segment that comprises a plurality of sector shaped abradable components 110, each having a support surface 112 retained within and coupled to the casing and an abradable substrate 120 that is in opposed, spaced relationship with the blade tip by a blade tip gap G. The support surface 112 has upstream and downstream ends relative to the turbine generalized flow direction F and a support surface axis that is parallel to the corresponding turbine blade rotational axis, which defines the curvature radius of the curved inwardly facing abradable substrate 120. Forward (upstream of the combustion hot working gas flow direction, F) and aft (downstream in the direction F) axial faces of the support surface 112 and the abradable substrate 120 are generally perpendicular to the support surface axis and the turbine blade rotational axis. The abradable substrate is often constructed of a metallic/ceramic material that has high thermal and thermal erosion resistance and that maintains structural integrity at high combustion temperatures. As the abradable surface 120 metallic ceramic materials is often more abrasive than the turbine blade tip 94 material a blade tip gap G is maintained to avoid contact between the two opposed components that might at best cause premature blade tip wear and in worse case circumstances might cause engine damage. Some known abradable components 110 are constructed with a monolithic metallic/ceramic abradable substrate 120. Other known abradable components 110 are constructed with a composite matrix composite (CMC) structure, comprising a ceramic support surface 112 to which is bonded a friable graded insulation (FGI) ceramic strata of multiple layers of closely-packed hollow ceramic spherical particles, surrounded by smaller particle ceramic filler, as described in U.S. Pat. No. 6,641,907. Spherical particles having different properties are layered in the substrate 120, with generally more easily abradable spheres forming the upper layer to reduce blade tip 94 wear. Another CMC structure is described in U.S. Patent Publication No. 2008/0274336, wherein the surface includes a cut-grooved pattern between the hollow ceramic spheres. The grooves are intended to reduce the abradable surface material cross sectional area to reduce potential blade tip 94 wear, if they contact the abradable surface. Other commonly known abradable components 110 are constructed with a metallic base layer support surface 112 to which is applied a thermally sprayed ceramic/metallic layer that forms the abradable substrate layer 120. As will be described in greater detail the thermally sprayed metallic layer may include grooves, depressions or ridges to reduce abradable surface material cross section for potential blade tip 94 wear reduction.
In addition to the desire to prevent blade tip 94 premature wear or contact with the abradable substrate 120, as shown in FIG. 3, for ideal airflow and power efficiency each respective blade tip 94 desirably has a uniform blade tip gap G relative to the abradable component 110 that is as small as possible (ideally zero clearance) to minimize blade tip airflow leakage L between the high pressure blade side 96 and the low pressure blade side 98 as well as axially in the combustion flow direction F. However, manufacturing and operational tradeoffs require blade tip gaps G greater than zero. Such tradeoffs include tolerance stacking of interacting components, so that a blade constructed on the higher end of acceptable radial length tolerance and an abradable component abradable substrate 120 constructed on the lower end of acceptable radial tolerance do not impact each other excessively during operation. Similarly, small mechanical alignment variances during engine assembly can cause local variations in the blade tip gap. For example in a turbine engine of many meters axial length, having a turbine casing abradable substrate 120 inner diameter of multiple meters, very small mechanical alignment variances can impart local blade tip gap G variances of a few millimeters.
During turbine engine 80 operation the turbine engine casing 100 may experience out of round (e.g., egg shaped) thermal distortion as shown in FIGS. 4 and 6. Casing 100 thermal distortion potential increases between operational cycles of the turbine engine 80 as the engine is fired up to generate power and subsequently cooled for servicing after thousands of hours of power generation. Commonly, as shown in FIG. 6, greater casing 100 and abradable component 110 distortion tends to occur at the uppermost 122 and lowermost 126 casing circumferential positions (i.e., 6:00 and 12:00 positions) compared to the lateral right 124 and left 128 circumferential positions (i.e., 3:00 and 9:00). If, for example as shown in FIG. 4 casing distortion at the 6:00 position causes blade tip contact with the abradable substrate 120 one or more of the blade tips may be worn during operation, increasing the blade tip gap locally in various other less deformed circumferential portions of the turbine casing 100 from the ideal gap G to a larger gap GW as shown in FIG. 5. The excessive blade gap Gw distortion increases blade tip leakage L, diverting hot combustion gas away from the turbine blade 92 airfoil, reducing the turbine engine's efficiency.
In the past flat abradable surface substrates 120 were utilized and the blade tip gap G specification conservatively chosen to provide at least a minimal overall clearance to prevent blade tip 94 and abradable surface substrate contact within a wide range of turbine component manufacturing tolerance stacking, assembly alignment variances, and thermal distortion. Thus, a relatively wide conservative gap G specification chosen to avoid tip/substrate contact sacrificed engine efficiency. Commercial desire to enhance engine efficiency for fuel conservation has driven smaller blade tip gap G specifications: preferably no more than 2 millimeters and desirably approaching 1 millimeter.
In order to reduce likelihood of blade tip/substrate contact, abradable components comprising metallic base layer supports with thermally sprayed metallic/ceramic abradable surfaces have been constructed with three dimensional plan form profiles, such as shown in FIGS. 7-11. The exemplary known abradable surface component 130 of FIGS. 7 and 10 has a metallic base layer support 131 for coupling to a turbine casing 100, upon which a thermally sprayed metallic/ceramic layer has been deposited and formed into three-dimensional ridge and groove profiles by known deposition or ablative material working methods. Specifically in these cited figures a plurality of ridges 132 respectively have a common height HR distal ridge tip surface 134 that defines the blade tip gap G between the blade tip 94 and it. Each ridge also has sidewalls 135 and 136 that extend from the substrate surface 137 and define grooves 138 between successive ridge opposed sidewalls. The ridges 132 are arrayed with parallel spacing SR between successive ridge centerlines and defined groove widths WG. Due to the abradable component surface symmetry, groove depths DG correspond to the ridge heights HR. Compared to a solid smooth surface abradable, the ridges 132 have smaller cross section and more limited abrasion contact in the event that the blade tip gap G becomes so small as to allow blade tip 94 to contact one or more tips 134. However, the relatively tall and widely spaced ridges 132 allow blade leakage L into the grooves 138 between ridges, as compared to the prior continuous flat abradable surfaces. In an effort to reduce blade tip leakage L, the ridges 132 and grooves 138 were oriented horizontally in the direction of combustion flow F (not shown) or diagonally across the width of the abradable surface 137, as shown in FIG. 7, so that they would tend to inhibit the leakage. Other known abradable components 140, shown in FIG. 8, have arrayed grooves 148 in crisscross patterns, forming diamond shaped ridge plan forms 142 with flat, equal height ridge tips 144. Additional known abradable components have employed triangular rounded or flat tipped triangular ridges 152 shown in FIGS. 9 and 11. In the abradable component 150 of FIGS. 9 and 11, each ridge 152 has symmetrical sidewalls 155, 156 that terminate in a flat ridge tip 154. All ridge tips 154 have a common height HR and project from the substrate surface 157. Grooves 158 are curved and have a similar plan form profile as the blade tip 94 camber line. Curved grooves 158 generally are more difficult to form than linear grooves 138 or 148 of the abradable components shown in FIGS. 7 and 8.
Past abradable component, designs have required stark compromises between blade tips wear resulting from contact between the blade tip and the abradable surface and blade tip leakage that reduces turbine engine operational efficiency. Optimizing engine operational efficiency required reduced blade tip gaps and smooth, consistently flat abradable surface topology to hinder air leakage through the blade tip gap, improving initial engine performance and energy conservation. In another drive for increased gas turbine operational efficiency and flexibility so-called “fast start” mode engines were being constructed that required faster full power ramp up (order of 40-50 MW/minute). Aggressive ramp-up rates exacerbated potential higher incursion of blade tips into ring segment abradable coating, resulting from quicker thermal and mechanical growth and higher distortion and greater mismatch in growth rates between rotating and stationary components. This in turn required greater turbine tip clearance in the “fast start” mode engines, to avoid premature blade tip wear, than the blade tip clearance required for engines that are configured only for “standard” starting cycles. Thus as a design choice one needed to balance the benefits of quicker startup/lower operational efficiency larger blade tip gaps or standard startup/higher operational efficiency smaller blade tip gaps.
Traditionally, standard or fast start engines required different construction to accommodate the different needed blade-tip gap parameters of both designs. Whether in standard or fast start configuration, decreasing blade tip gap for engine efficiency optimization ultimately risked premature blade tip wear, opening the blade tip gap and ultimately decreasing longer-term engine performance efficiency during the engine operational cycle. The aforementioned ceramic matrix composite (CMC) abradable component designs sought to maintain airflow control benefits and small blade tip gaps of flat surface profile abradable surfaces by using a softer top abradable layer to mitigate blade tip wear. The abradable components of the U.S. Patent Publication No. 2008/0274336 also sought to reduce blade tip wear by incorporating grooves between the upper layer hollow ceramic spheres. However, groove dimensions were inherently limited by the packing spacing and diameter of the spheres in order to prevent sphere breakage. Adding uniform height abradable surface ridges to thermally sprayed substrate profiles as a compromise solution to reduce blade tip gap while reducing potential rubbing contact surface area between the ridge tips and blade tips reduced likelihood of premature blade tip wear/increasing blade tip gap but at the cost of increased blade tip leakage into grooves between ridges. As noted above, attempts have been made to reduce blade-tip leakage flow by changing plan form orientation of the ridge arrays to attempt to block or otherwise control leakage airflow into the grooves.